Gas turbine engine airfoil platform cooling

ABSTRACT

An airfoil for a gas turbine engine includes an airfoil that extends from a platform that has first and second circumferential sides that respectively extend to first and second circumferential edges. The first circumferential side has a tapered surface at a first angle relative to a flow path surface. The second circumferential surface has a cooling hole that extends toward the second lateral edge at a second angle relative to the flow path surface. The tapered surface and the cooling hole are axially aligned with one another.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. Provisional Application No.62/008,599 which was filed on Jun. 6, 2014 and is incorporated herein byreference.

BACKGROUND

This disclosure relates to a gas turbine engine airfoil, and moreparticularly, a platform cooling arrangement.

Industrial gas turbine engines include a compressor section, a combustorsection and a turbine section. Air entering the compressor section iscompressed and delivered into the combustor section where it is mixedwith fuel and ignited to generate a high-speed exhaust gas flow. Thehigh-speed exhaust gas flow expands through the turbine section to drivethe compressor and a power turbine. The compressor and turbine sectionseach included multiple circumferential arrays of blades and vanes.

The turbine section in particular is subject to high temperatures thatmay exceed the melting temperature of the components. To this end, thesecomponents are cooled by one or more cooling mechanisms. Airfoils extendfrom a platform, and in the case of a blade, an inner platform supportedby a root section. The airfoil and platform typically included coolingholes to supply cooling fluid to the hotter areas of the blade.

As gas turbine engines are pushed to higher temperatures to increasepower output and efficiency, distress of the airfoil platformsincreasingly becomes the service life limiting area. A typical solutionto this includes decreasing the platform metal temperatures with coolingair. One approach includes providing platform cooling holes suppliedwith “wheel-space air,” which corresponds to fluid provided betweenadjacent turbine blade shanks. This approach may supply insufficientpressure needed for adequate film cooling to the local and adjacentplatform. Another approach includes providing cooling to the platformmate faces, or facing edges, supplied by cooling air from the airfoilcore and/or “wheel-space air.” This approach may not provide filmcooling and subjects the area to cracking due to reduced wall thickness.

SUMMARY

In one exemplary embodiment, an airfoil for a gas turbine engineincludes an airfoil that extends from a platform that has first andsecond circumferential sides that respectively extend to first andsecond circumferential edges. The first circumferential side has atapered surface at a first angle relative to a flow path surface. Thesecond circumferential surface has a cooling hole that extends towardthe second lateral edge at a second angle relative to the flow pathsurface. The tapered surface and the cooling hole are axially alignedwith one another.

In a further embodiment of the above, the airfoil includes a coolingpassage. The cooling hole is in fluid communication with the coolinghole.

In a further embodiment of any of the above, the second angle is 5-40°.

In a further embodiment of any of the above, the second angle is 15-30°.

In a further embodiment of any of the above, multiple cooling holes arearranged in a cluster. The cluster is arranged near a trailing edge ofthe airfoil on a pressure side.

In a further embodiment of any of the above, the cluster is arrangedwithin about three inches (76.2 mm) of an aft edge of the platform.

In a further embodiment of any of the above, the cluster is within about0.6 inch (15.2 mm) of the second lateral edge.

In a further embodiment of any of the above, the cooling holes each havea diameter equivalent of 0.010-0.050 inch (0.25-1.27 mm).

In a further embodiment of any of the above, the first angle is 1-20°.

In a further embodiment of any of the above, the first angle is 2-15°.

In a further embodiment of any of the above, the tapered surface extendswithin about three inches (76.2 mm) of the first circumferential edge toan aft edge of the platform.

In a further embodiment of any of the above, the tapered surface extendsless that 0.7 inch (17.78 mm) from the first circumferential edge.

In a further embodiment of any of the above, the airfoil is a turbineblade.

In another exemplary embodiment, an array of airfoils for a gas turbineengine includes adjacent airfoils. Each airfoil extends from a platformthat has first and second circumferential sides respectively that extendto first and second circumferential edges. The first circumferentialside has a tapered surface at a first angle relative to a flow pathsurface. The second circumferential surface has a cooling hole thatextends toward the second lateral edge at a second angle relative to theflow path surface. The tapered surface and the cooling hole are axiallyaligned with one another.

In a further embodiment of the above, the airfoils include a coolingpassage. The cooling hole is in fluid communication with the coolingpassage.

In a further embodiment of any of the above, the second angle is 5-40°and comprises multiple cooling holes that are arranged in a cluster. Thecluster is arranged near a trailing edge of the airfoil on a pressureside.

In a further embodiment of any of the above, the cluster is arrangedwithin about three inches (76.2 mm) of an aft edge of the platform. Thecluster is within about 0.6 inch (15.2 mm) of the second lateral edge.The cooling holes each have a diameter equivalent of 0.010-0.050 inch(0.25-1.27 mm).

In a further embodiment of any of the above, the first angle is 1-20°.

In a further embodiment of any of the above, the tapered surface extendswithin about three inches (76.2 mm) of the first circumferential edge toan aft edge of the platform. The tapered surface extends less that 0.7inch (17.78 mm) from the first circumferential edge.

In a further embodiment of any of the above, the airfoil is a turbineblade.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 is a schematic cross-sectional view of an example industrial gasturbine engine.

FIG. 2 schematically illustrates a section of the gas turbine engine,such as a turbine section.

FIGS. 3A and 3B are perspective and elevational views respectively ofadjacent blades.

FIG. 4 is an elevational view of the blade shown in FIGS. 3A and 3B.

FIG. 5 is an enlarged cross-sectional view of the adjacent blades.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

A schematic view of an industrial gas turbine engine 10 is illustratedin FIG. 1. The engine 10 includes a compressor section 12 and a turbinesection 14 interconnected to one another by a shaft 16 rotatable aboutan axis X. A combustor 18 is arranged between the compressor and turbinesections 12, 14. A generator 22 is rotationally driven by a shaftcoupled to the turbine or uncoupled via a power turbine 20, which isconnected to a power grid 23. It should be understood that theillustrated engine 10 is highly schematic, and may vary from theconfiguration illustrated. Moreover, the disclosed airfoil may be usedin commercial and military aircraft engines as well as industrial gasturbine engines.

The turbine section 14 includes multiple turbine blades, one of which isillustrated at 64 in FIG. 2. In the example turbine section 14, firstand second arrays of circumferentially spaced fixed vanes 60, 62 areaxially spaced apart from one another. A first stage array ofcircumferentially spaced turbine blades 64, mounted to a rotor disk 68,is arranged axially between the first and second fixed vane arrays. Asecond stage array of circumferentially spaced turbine blades 66 isarranged aft of the second array of fixed vanes 62. It should beunderstood that any number of stages may be used. Moreover, thedisclosed airfoil may be used in a compressor section, turbine sectionand/or fixed or rotating stages.

The turbine blades each include a tip 80 adjacent to a blade outer airseal 70 of a case structure 72, which provides an outer flow path. Thefirst and second stage arrays of turbine vanes and first and secondstage arrays of turbine blades are arranged within a core flow path Cand are operatively connected to the shaft 16, for example.

Each blade 64 includes an inner platform 76 respectively defining aninner flow path. The platform inner platform 76 supports an airfoil 78that extends in a radial direction R. It should be understood that theturbine blades may be discrete from one another or arranged inintegrated clusters. The airfoil 78 provides leading and trailing edges82, 84.

The airfoil 78 is provided between pressure (typically concave) andsuction (typically convex) sides in circumferential direction Y (FIG. 4)provided between the leading and trailing edges 82, 84. The turbineblades 64 are constructed from a high strength, heat resistant materialsuch as a nickel-based or cobalt-based superalloy, or of a hightemperature, stress resistant ceramic or composite material. In cooledconfigurations, internal fluid passages and external cooling aperturesprovide for a combination of impingement and film cooling. Other coolingapproaches may be used such as trip strips, pedestals or otherconvective cooling techniques. In addition, one or more thermal barriercoatings, abrasion-resistant coatings or other protective coatings maybe applied to the turbine vane 64.

The airfoil 78 extends from the platform 76 and provides first andsecond circumferential sides, which corresponds to pressure and suctionsides 86, 88, as shown in Figure 4. With continued reference to FIG. 4,the first and second circumferential sides 86, 88 include first andsecond circumferential edges 92, 94, respectively. This firstcircumferential side 86 has a tapered surface 90 at a first angle 100relative to the flowpath surface provided by the platform 76. The firstangle is 1-20 degrees, and in another example, 2-15 degrees. In stillanother example, the first angle is 2-12 degrees.

The tapered surface 90 extends within about three inches (76.2 mm) ofthe first circumferential edge 92 to an aft edge 120 of the platform 76.The tapered surface 90 extends a width 108 less that 0.7 inch (17.78 mm)from the first edge 92.

The second circumferential surface 88 has at least one cooling hole 98,for example, a cluster of cooling holes, extending toward the secondcircumferential edge 92 at a second angle 96 relative to the flowpathsurface. As shown in FIG. 5, the airfoil 78 includes a cooling passage99 in fluid communication with the cooling hole 98. In one example, thesecond angle 96 is 5-40 degrees, in another example, the second angle 96is 15-30 degrees.

The tapered surface 90 and the cooling hole 98 are axially aligned withone another such that cooling fluid from the cooling hole 98 is directedtoward the tapered surface 90. The cooling arrangement provides for amore effective platform cooling. The relationship between these featuresand adjacent blades is shown in FIGS. 3A-3B.

The cluster of cooling holes 98 is arranged within about 3 inches (76.2mm) of an aft edge 120 of the platform 76. The cluster is within about0.60 inch (15.2 mm) of the second lateral edge 90. A lateral offset 110is about 0.60 inch (15.2 mm), and a lateral width 112 is at least 0.010inch (0.25 mm). Each cooling hole 98 has an effective diameter, ordiameter equivalent of 0.010-0.050 inch (0.25 -1.27 mm). The clusterincludes an axial offset 114 from the aft edge 120 of about 1 inch (25.4mm) and may extend a distance 116, which is less than three inches (76.2mm). The cooling holes may be round or shaped. The holes may have auniform cross-section or may be shaped as diffusers.

A seal 118 may be provided between adjacent blades 64 to obstruct thegap provided between the facing circumferential edges 92, 94. The seal118 is schematically illustrated and may be provided using any suitablearrangement.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. An airfoil for a gas turbine engine comprising:an airfoil extending from a platform that has first and secondcircumferential sides respectively extending to first and secondcircumferential edges, the first circumferential side having a taperedsurface at a first angle relative to a flow path surface, and the secondcircumferential surface having a cooling hole extending toward thesecond lateral edge at a second angle relative to the flow path surface,the tapered surface and the cooling hole axially aligned with oneanother.
 2. The airfoil according to claim 1, wherein the airfoilincludes a cooling passage, the cooling hole in fluid communication withthe cooling hole.
 3. The airfoil according to claim 2, wherein thesecond angle is 5-40°.
 4. The airfoil according to claim 3, wherein thesecond angle is 15-30°.
 5. The airfoil according to claim 2, comprisingmultiple cooling holes arranged in a cluster, the cluster arranged neara trailing edge of the airfoil on a pressure side.
 6. The airfoilaccording to claim 5, wherein the cluster is arranged within about threeinches (76.2 mm) of an aft edge of the platform.
 7. The airfoilaccording to claim 5, wherein the cluster is within about 0.6 inch (15.2mm) of the second lateral edge.
 8. The airfoil according to claim 5,wherein the cooling holes each have a diameter equivalent of 0.010-0.050inch (0.25-1.27 mm).
 9. The airfoil according to claim 1, wherein thefirst angle is 1-20°.
 10. The airfoil according to claim 9, wherein thefirst angle is 2-15°.
 11. The airfoil according to claim 9, wherein thetapered surface extends within about three inches (76.2 mm) of the firstcircumferential edge to an aft edge of the platform.
 12. The airfoilaccording to claim 11, wherein the tapered surface extends less that 0.7inch (17.78 mm) from the first circumferential edge.
 13. The airfoilaccording to claim 1, wherein the airfoil is a turbine blade.
 14. Anarray of airfoils for a gas turbine engine comprising: adjacentairfoils, each airfoil extending from a platform that has first andsecond circumferential sides respectively extending to first and secondcircumferential edges, the first circumferential side having a taperedsurface at a first angle relative to a flow path surface, and the secondcircumferential surface having a cooling hole extending toward thesecond lateral edge at a second angle relative to the flow path surface,the tapered surface and the cooling hole axially aligned with oneanother.
 15. The array of airfoils according to claim 14, wherein theairfoils includes a cooling passage, the cooling hole in fluidcommunication with the cooling passage.
 16. The array of airfoilsaccording to claim 15, wherein the second angle is 5-40°, and comprisingmultiple cooling holes arranged in a cluster, the cluster arranged neara trailing edge of the airfoil on a pressure side.
 17. The array ofairfoils according to claim 16, wherein the cluster is arranged withinabout three inches (76.2 mm) of an aft edge of the platform, the clusteris within about 0.6 inch (15.2 mm) of the second lateral edge, whereinthe cooling holes each have a diameter equivalent of 0.010-0.050 inch(0.25-1.27 mm).
 18. The array of airfoils according to claim 16, whereinthe first angle is 1-20°.
 19. The array of airfoils according to claim18, wherein the tapered surface extends within about three inches (76.2mm) of the first circumferential edge to an aft edge of the platform,wherein the tapered surface extends less that 0.7 inch (17.78 mm) fromthe first circumferential edge.
 20. The array of airfoils according toclaim 14, wherein the airfoil is a turbine blade.